Gas turbine engines are typically employed to power aircraft. Typically a gas turbine engine will comprise an axial fan driven by an engine core. The engine core is generally made up of one or more turbines which drive respective compressors via coaxial shafts. The fan is usually driven off an additional lower pressure turbine in the engine core.
A compressor typically comprises a series of rotatable components. The rotatable components each comprise an array of blades, each having an aerofoil cross section. The blades are attached to a central hub or drum. The blades of the rotatable components accelerate the air through the engine. Each of the rotatable components are coupled with a static component which comprises an array of vanes that are also of aerofoil cross section. The static components are connected to radially inner and/or outer casing components. Conventionally, the surface of the hub and the casing components that define the main gas flow path through the compressor are referred to as endwalls.
Compressors are designed to avoid or limit stall. Stall is a disruption of airflow through the compressor that can result in a momentary power drop or in the worst case a complete loss of compression. It is desirable to increase the range at which the compressor can operate before experiencing stall (the stall margin).
Near the endwalls of blade rows in a multi-stage compressor, a spanwise region of low axial momentum flow develops. This region of low axial momentum flow leads to two flow regimes: that of the endwall region and that of the free-stream. The endwall region is influenced by the presence of the hub and casing components, clearance flows, and corner separations, and has a high degree of three dimensional flow. The free-stream can be treated as a flow region unaffected by endwall effects. The extent of the endwall and freestream regimes vary depending on aerodynamic loading. The further off-design (i.e. the closer to stall), the larger the three dimensional flow regime becomes.
Conventionally, gas turbine engine compressor designers have attempted to increase the stall margin by delaying the rise in endwall loss. A common way to increase stall margin is to reduce the pitch-to-chord ratio of the blade row, which can be achieved by either increasing blade count and/or increasing the chord. Having a smaller pitch-to-chord delays the growth of the corner separation. The number of blades and chord is therefore strongly linked to the stall margin requirement.
However, the number of blades required for sufficient stall margin is often greater than the optimal number for best design point efficiency. As such, designers attempt to delay the growth of the corner separation by means other than increasing blade count or chord. Typically, the stall margin can be increased by delaying the growth in corner separation by leaning blades and/or vanes and changing the sweep of the blades and/or vanes. Attempts have also been made to increase the stall margin by de-cambering blades (i.e. reducing the curvature of the blades) in a region near to the hub and casing component.